Preferential flow distribution for gas turbine engine component

ABSTRACT

A combustor liner for a gas turbine engine according to an example of the present disclosure includes, among other things, at least one liner segment that has an external wall dimensioned to bound a combustion chamber. The external wall extends between leading and trailing edges in an axial direction and extends between opposed mate faces in a circumferential direction. A cooling circuit is defined by the external wall. A plurality of heat transfer features are distributed in the cooling circuit to define a first restricted flow region that tapers from the leading edge to the trailing edge and to define at least one prioritized flow region that extends substantially from the leading edge to the trailing edge such that the at least one prioritized flow region is bounded by a perimeter of the first restricted flow region, and the at least one prioritized flow region has a lesser concentration of the plurality of heat transfer features than the first restricted flow region.

BACKGROUND

This disclosure relates to a combustor for a gas turbine engine and,more particularly, to flow distribution through a combustor liner of thecombustor.

Gas turbine engines can include a fan for propulsion air and to coolcomponents. The fan also delivers air into a core engine where it iscompressed. The compressed air is then delivered into a combustorsection.

The combustor section includes one or more combustor liners that definea combustion chamber. Fuel is ejected from fuel injectors into thecombustion chamber. The compressed air is mixed with the fuel andignited in the combustion chamber to produce relatively hot combustiongases. The combustion gases expand downstream over and drive turbineblades.

The combustor liners are subject to extreme heat due to the combustionprocess. Formation of hot spots can occur along localized regions of thecombustor liners. Cooling flow may be utilized to cool portions of thecombustor liners at locations adjacent to the hot spots.

SUMMARY

A combustor liner for a gas turbine engine according to an example ofthe present disclosure includes at least one liner segment that has anexternal wall dimensioned to bound a combustion chamber. The externalwall extends between leading and trailing edges in an axial directionand extends between opposed mate faces in a circumferential direction. Acooling circuit is defined by the external wall. A plurality of heattransfer features are distributed in the cooling circuit to define afirst restricted flow region that tapers from the leading edge to thetrailing edge and to define at least one prioritized flow region thatextends substantially from the leading edge to the trailing edge suchthat the at least one prioritized flow region is bounded by a perimeterof the first restricted flow region, and the at least one prioritizedflow region has a lesser concentration of the plurality of heat transferfeatures than the first restricted flow region.

In a further embodiment of any of the foregoing embodiments, the atleast one prioritized flow region includes first and second prioritizedflow regions on opposed sides of the first restricted flow region.

In a further embodiment of any of the foregoing embodiments, theplurality of heat transfer features are distributed in the coolingcircuit to define second and third restricted flow regions that extendsubstantially along the mate faces to bound respective ones of the firstand second prioritized flow regions.

In a further embodiment of any of the foregoing embodiments, each of thefirst and second prioritized flow regions has a substantiallytrapezoidal geometry.

In a further embodiment of any of the foregoing embodiments, the atleast one liner segment includes a first liner segment and a secondliner segment arranged in the axial direction to define a stepwisechange in area of the combustion chamber such that the cooling circuitof the first liner segment is oriented to eject cooling flow from thetrailing edge of the first liner segment onto external surfaces of theexternal wall of the second liner segment that defines the combustionchamber.

In a further embodiment of any of the foregoing embodiments, the atleast one liner segment includes an array of liner segments, and each ofthe mate faces defines an intersegment gap with an adjacent one of theliner segments.

In a further embodiment of any of the foregoing embodiments, theplurality of heat transfer features includes a plurality of pedestalsthat extend in a radial direction between opposed internal surfacesdefining the cooling circuit.

In a further embodiment of any of the foregoing embodiments, respectivesets of the plurality of heat transfer features are uniformlydistributed in the first restricted flow region and in the at least oneprioritized flow region.

A further embodiment of any of the foregoing embodiments includes athermal barrier coating disposed on surfaces of the external walldefining the combustion chamber.

In a further embodiment of any of the foregoing embodiments, thesurfaces of the external wall defining the combustion chamber aresubstantially free of any cooling apertures along the cooling circuit.

In a further embodiment of any of the foregoing embodiments, theexternal wall is a bulkhead that bounds the combustion chamber in theaxial direction. The bulkhead has at least one aperture along thecombustion chamber that is dimensioned to receive a fuel injectornozzle.

A combustor section for a gas turbine engine according to an example ofthe present disclosure includes an array of fuel injector nozzlesarranged about a longitudinal axis. A combustor liner includes an arrayof liner segments arranged about the longitudinal axis to define acombustion chamber. Each one of the fuel injector nozzles defines anozzle axis. A projection of the nozzle axis extends through thecombustion chamber. Each one of the liner segments includes an externalwall extending axially between leading and trailing edges and extendingcircumferentially between opposed mate faces with respect to thelongitudinal axis. A cooling circuit is defined by the external wall. Aplurality of heat transfer features are distributed in the coolingcircuit to define first and second prioritized flow regions on opposedsides of a first restricted flow region. Each of the first and secondprioritized flow regions extend axially along the projection of thenozzle axis of respective ones of the fuel injector nozzles from theleading edge to the trailing edge such that the first restricted flowregion tapers from the leading edge to the trailing edge, and each ofthe first and second prioritized flow regions have a relatively greateraverage flow path volume than the first restricted flow region.

In a further embodiment of any of the foregoing embodiments, theplurality of heat transfer features are distributed in the coolingcircuit to define second and third restricted flow regions that extendsubstantially along the mate faces to bound a perimeter of respectiveones of the first and second prioritized flow regions, and each of thefirst and second prioritized flow regions has a relatively greateraverage flow path volume than the second and third restricted flowregions.

In a further embodiment of any of the foregoing embodiments, surfaces ofthe external wall defining the combustion chamber are substantially freeof any cooling apertures along the cooling circuit.

A gas turbine engine according to an example of the present disclosureincludes a compressor section, a turbine section that drives thecompressor section, a combustor section has a combustor. The combustorhas a combustor liner and an array of fuel injector nozzles arrangedabout an engine longitudinal axis. The combustor liner has an array ofliner segments arranged about the engine longitudinal axis to define acombustion chamber. Each one of the fuel injector nozzles defines anozzle axis. A projection of the nozzle axis extends through thecombustion chamber. Each one of the liner segments includes an externalwall extending axially between leading and trailing edges and extendingcircumferentially between opposed mate faces with respect to the enginelongitudinal axis. A cooling circuit is defined by the external wall. Aplurality of heat transfer features are distributed in the coolingcircuit to define first and second prioritized flow regions on opposedsides of a first restricted flow region. Each of the first and secondprioritized flow regions extend axially along the projection of thenozzle axis of respective ones of the fuel injector nozzles such that awidth of the first and second prioritized flow regions progressivelyincreases from the leading edge to the trailing edge.

In a further embodiment of any of the foregoing embodiments, the firstrestricted flow region is circumferentially spaced from the projectionof the nozzle axis of each and every one of the fuel injector nozzles.

In a further embodiment of any of the foregoing embodiments, the firstrestricted flow region tapers from the leading edge to the trailingedge.

In a further embodiment of any of the foregoing embodiments each of themate faces defines an intersegment gap with the mate face of an adjacentone of the liner segments. The plurality of heat transfer features aredistributed in the cooling circuit to define second and third restrictedflow regions that extend substantially along the mate faces to bound aperimeter of respective ones of the first and second prioritized flowregions.

In a further embodiment of any of the foregoing embodiments, theintersegment gap is dimensioned to eject cooling flow into thecombustion chamber.

In a further embodiment of any of the foregoing embodiments, the arrayof liner segments includes a first set of liner segments and a secondset of liner segments axially arranged to define a stepwise change inarea of the combustion chamber such that each cooling circuit of thefirst set of liner segments is oriented to eject cooling flow ontoexternal surfaces of the second set of liner segments bounding thecombustion chamber.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of an embodiment. The drawings that accompany the detaileddescription can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 is a perspective view of a combustor.

FIG. 3 is a sectional view of the combustor taken along line 3-3 of FIG.2.

FIG. 3A schematically illustrates the combustor taken along line 3A-3Aof FIG. 3.

FIG. 4 is a sectional view of the combustor of FIG. 3 with a fuelinjector nozzle uninstalled.

FIG. 5 is a sectional view of adjacent liner segments of the combustorof FIG. 3.

FIG. 6 is a sectional view of the liner segments taken along line 6-6 ofFIG. 5.

FIG. 7 illustrates an example distribution of heat transfer features.

FIG. 8 is a sectional view of a liner segment according to anotherexample.

FIG. 9 is a sectional view of a liner segment arranged relative to afuel injector nozzle according to yet another example.

FIG. 10 is a sectional view of a liner segment arranged relative to afuel injector nozzle according to yet another example.

FIG. 11 illustrates example geometries of heat transfer features.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle15, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

Referring to FIG. 2, the combustor 56 includes at least one combustorcase 58 that extends along a longitudinal axis X. The longitudinal axisX can be parallel to or collinear with the engine longitudinal axis A ofFIG. 1. The combustor case 58 includes an inner (or first) combustorcase 58A and an outer (or second) diffuser case 58B that extend aboutthe longitudinal X. Each of the cases 58A, 58B has a generally annulargeometry.

Referring to FIG. 3, with continuing reference to FIG. 2, a divergentnozzle or diffuser 55 is dimensioned to deliver flow in the core flowpath C from the compressor section 24 (FIG. 1) to the combustor case 58.The combustor 56 includes a plurality of combustor liners 60 arrangedbetween the cases 58A, 58B. The combustor liners 60 include at least aninner (or first) combustor liner 60A and an outer (or second) combustorliner 60B that are concentric and are arranged to extend about thelongitudinal axis X.

The inner combustion liner 60A extends about the inner combustor case58A to define an inner (or first) plenum 62. The outer diffuser case 58Bextends about the outer combustor liner 60B to define an outer (orsecond) plenum 63. Each of the plenums 62, 63 has a generally annulargeometry. The plenums 62, 63 can be arranged to receive flow from thediffusor 55.

Each combustor liner 60 can include one or more liner segments 68. Theliner segments 68 have an arcuate geometry and are arranged in an arrayabout the longitudinal axis X to bound or otherwise define an annularcombustion chamber 64, as illustrated schematically by FIG. 3A. Theliner segments 68 can be made of a high temperature metal or metalalloy, including directionally solidified and single crystal materials,for example.

The combustor 56 includes a bulkhead 60C that bounds the combustionchamber 64 in an axial direction with respect to the longitudinal axisX. The combustor 56 includes an array of fuel injectors 66 arrangedabout the longitudinal axis X, as illustrated by FIGS. 3 and 3A. Eachfuel injector 66 is fluidly coupled to a fuel source FS. The fuel sourceFS is operable to supply fuel to each fuel injector 66 during engineoperation.

Each fuel injector 66 includes a fuel injector nozzle 67 that isoperable to eject a quantity of fuel FF along a respective nozzle axisN. A projection of the nozzle axis N extends through the combustionchamber 64. A major component of the nozzle axis N extends in adirection that is parallel to the longitudinal axis X, as illustrated byFIG. 3.

Referring to FIG. 4, with continuing reference to FIG. 3, the combustorliners 60 are shown with the fuel injector nozzle 67 of FIG. 3 removedfor illustrative purposes. An injector mount 65 is dimensioned toreceive a respective one of the nozzles 67 and extends along arespective nozzle axis N. The bulkhead 60C can define one or moreapertures 61 (one shown for illustrative purposes) defined along thecombustion chamber 64. Each aperture 61 is dimensioned to receive arespective fuel injector nozzle 67, as illustrated by FIG. 3.

Each combustor liner 60A, 60B can include a liner support 69 thatextends in the axial direction from the bulkhead 60C. Each liner segment68 can be mounted or otherwise mechanically attached to the linersupport 69 with one or more fasteners 73, for example.

The liner support 69 can have a stepwise geometry, with the linersegments 68 arranged in the axial direction with respect to thelongitudinal axis X to define a stepwise change in area of thecombustion chamber 64 along the liner support 69. Adjacent linersegments 68 can axially overlap relative to the longitudinal axis X.

Each liner segment 68 defines a cooling circuit 70 that conveys coolingflow CF to cool portions of the liner segment 68 and adjacent portionsof the combustor 56, such as an adjacent (e.g., downstream) linersegment 68. As illustrated by FIGS. 4 and 5, the array of liner segments68 can be arranged in the axial direction with respect to longitudinalaxis X to define a step formation or stepwise change in area of thecombustion chamber 64 such that each cooling circuit 70 of an upstream(or first) set of liner segments 68 is oriented to eject cooling flow CFfrom a respective trailing edge 68TE onto external surfaces of eachexternal wall 68A of a downstream (or second) set of liner segments 68bounding the combustion chamber 64, as illustrated by liner segments68-1, 68-2.

FIG. 6 illustrates a sectional view of one of the liner segments 68defining a respective cooling circuit 70. The external wall 68A of theliner segment 68 is dimensioned to bound the combustion chamber 64. Theexternal wall 68A extends between leading and trailing edges 68LE, 68TEin the axial direction and extends between opposing mate faces 68M in acircumferential direction with respect to the longitudinal axis X. Eachof the liner segments 68 can have a substantially rectangularcross-sectional geometry.

In examples, the external wall 68A is a portion of one of the innerand/or outer combustor liners 60A, 60B or bulkhead 60C (FIGS. 3 and 4).For example, the liner segment 68 can be one of the liner segments 68 ofthe inner combustion liner 60A. In the illustrative example of FIG. 6,the liner segment 68 is the upstream liner segment 68-1 of outercombustion liner 60B positioned axially forward of axially aft linersegment 68-2 (shown in dashed lines for illustrative purposes). Linersegment 68-1 can be situated in a first, axially forwardmost row of theliner segments 68 relative to the fuel injector nozzles 67 and bulkhead60C, as illustrated by FIGS. 3 and 4, or can be situated in another oneof the rows of liner segments 68 such as the second row of linersegments 68-2. Although the cooling circuits disclosed herein primarilyrefer to a combustor, other gas turbine engine components such as wallsof the core flow path C in the turbine section 28 or mid-turbine frame57 and other systems requiring cooling augmentation can benefit from theteachings disclosed herein.

The cooling circuit 70 is defined by surfaces of the external wall 68A.In the illustrated example of FIG. 5, the combustion chamber 64 and thecooling circuit 70 are on opposed sides of the external wall 68A. Insome examples, a thermal barrier coating 71 is disposed on surfaces ofthe external wall 68A (shown in dashed lines in FIG. 5 for illustratedpurposes). In the illustrative example of FIGS. 5 and 6, surfaces of theexternal wall 68A defining the combustion chamber 64 are substantiallyfree of any cooling apertures along the cooling circuit 70.

The liner segment 68 is circumferentially aligned with the nozzle axes Nof two fuel injector nozzles 67. The projections of the nozzle axes Ncan be relatively closer to the mate faces 68M than arrangements havinga liner segment circumferentially aligned with the nozzle axis of onlyone fuel injector nozzle, which may cause relatively greater thermalgradients to form across the liner segment 68 and non-uniformdistribution of heat. The thermal gradients may cause the liner segment68 to expand and distort during engine operation. The cooling circuitsdisclosed herein can be arranged to reduce the formation of thermalgradients across the liner segments.

Liner segment 68 includes a plurality of heat transfer features 72extending from the external wall 68A. The heat transfer features 72 aredistributed in the cooling circuit 70 to interact with cooling flow CFfor providing convective cooling to adjacent portions of the linersegment 68. The heat transfer features 72 extend in a radial directionwith respect to the longitudinal axis X at least partially betweenopposed internal surfaces of the external wall 68A and liner support 69that define the cooling circuit 70. In the illustrative example of FIG.5, heat transfer features 72 include pin-fins or pedestals that extendin the radial direction between the opposed internal surfaces of theexternal wall 68A and liner support 69 that define the cooling circuit70. Each of the pedestals can have an elliptical geometry, asillustrated by FIG. 6. However, other geometries can be utilized such asrectangular, cube, diamond, oblong, teardrop, triangular or racetrackshaped cross-sectional geometries. One would understand how to dimensionthe heat transfer features 72 utilizing the teachings disclosed herein.

The heat transfer features 72 are arranged in contiguous sets 72-1through 72-3. Each of the respective sets 72-1 through 72-3 of heattransfer features 72 can be uniformly distributed, as illustrated byFIG. 6. In other examples, at least some of the heat transfer features72′ in the respective sets 72-1′, 72-2′ and/or 72-3′ are non-uniformlydistributed, as illustrated by FIG. 7.

Each of the sets of heat transfer features 72-1, 72-2, 72-3 can bearranged relative to non-uniform boundary conditions such as heatconcentrations or localized hotspots HS (shown in dashed lines in FIG. 6for illustrative purposes) that can form along the liner segment 68during engine operation. For example, formation of localized hot spotsHS can occur due to ignition of fuel FF ejected by the nozzles 67. Thehot spots HS can be generated or otherwise formed along the respectivenozzle axes N. The hot spots HS typically have a relatively greatertemperature than other portions of the liner segment 68 in operation,and in some scenarios can establish a peak temperature gradient relativeto the liner segment 68. The distribution of heat transfer features 72can reduce a likelihood of degradation of the liner segment 68 adjacentthe hot spots HS that may otherwise occur due to excessive temperatureexposure and insufficient cooling augmentation.

The sets of heat transfer features 72-1 to 72-3 are distributed in thecooling circuit 70 to define at least one prioritized flow region and atleast one restricted flow region to prioritize distribution of coolingflow CF in the cooling circuit 70. In the illustrated example of FIG. 6,the cooling circuit 70 includes first and second prioritized flowregions 74-1, 74-2 that extend along and are defined on opposed sides ofa first restricted flow region 74-3.

The flow regions 74 can be dimensioned with respect to the location ofeach of the nozzle axes N. For example, the heat transfer features 72are arranged such that the prioritized flow regions 74-1, 74-2 extendalong the projection of a respective one of the nozzle axes N, and therestricted flow region 74-3 is circumferentially spaced from theprojection of each and every one of the nozzle axes N with respect tothe longitudinal axis X.

A concentration of heat transfer features 72 in each flow region 74 canbe defined with respect to a volume of the cooling circuit 70 per unitarea, which can be set by the shape, spacing, size and/or orientation ofthe heat transfer features 72. Each of the prioritized flow regions74-1, 74-2 has a relatively lesser concentration of the heat transferfeatures 72 than the restricted flow region 74-3. An averageconcentration of heat transfer features 72 in the restricted flow region74-3 differs in the circumferential direction from the prioritized flowregions 74-1, 74-2 for at least a majority of axial positions relativeto the longitudinal axis X.

In the illustrative example of FIG. 6, the heat transfer features 72-3in the restricted flow region 74-3 are more densely spaced than the heattransfer features 72-1, 72-2 in the prioritized flow regions 74-1, 74-2.The relative concentrations of the heat transfer features 72-1, 72-2,72-3 can increase the amount of convective cooling to portions of theliner segment 68 adjacent the localized hot spots HS, even though theconcentration of heat transfer features 72-1, 72-2 in the prioritizedflow regions 74-1, 74-2 is less than the concentration of heat transferfeatures 72-3 in the restricted flow region 74-3.

The sets of heat transfer features 72-1, 72-2, 72-3 establish respectiveperimeters P1, P2, P3 (shown in dashed lines) of the prioritized andrestricted flow regions 74-1, 74-2, 74-3. The prioritized flow regions74-1, 74-2 extend substantially from the leading edge 68LE to thetrailing edge 68TE such that the perimeters P1, P2 are bounded by theperimeter P3 of the restricted flow region 74-3. The perimeters P1, P2of the prioritized flow regions 74-1, 74-2 extend substantially along arespective one of the mate faces 68M. For the purposes of thisdisclosure, the term “substantially” means that the respective perimeterP1/P2/P3 is defined within an average distance of the respective heattransfer features 72-1/72-2/72-3 from the referenced component, such asthe leading edge 68LE, mate faces 68M and/or trailing edge 68TE.

The heat transfer features 72 can be dimensioned and arranged such thateach of the prioritized flow regions 74-1, 74-2 has a relatively greateraverage flow path volume than the restricted flow region 74-3. Theaverage flow path volume can be defined as a volume of the coolingcircuit 70 within the respective perimeter P1, P2, P3 per unit area.

In examples, each of the prioritized flow regions 74-1, 74-2 comprisesat least 25% of a total flow path volume of the cooling circuit 70 inthe liner segment 68, or more narrowly between 30% and 40% of the totalflow path volume, with the restricted flow region 74-3 comprising aremainder of the total flow path volume. In examples, the prioritizedflow regions 74-1, 74-2 have at least a quantity of three or four heattransfer features 72 per square inch for at least a majority of thecross sectional area of the respective prioritized flow regions 74-1,74-2.

The flow regions 74 can be dimensioned relative to the localized hotspots HS. In the illustrative example of FIG. 6, the perimeter P3 of therestricted flow region 74-3 has a substantially trapezoidal (e.g.,isosceles) geometry, with a width of the restricted flow region 74-3generally increasing from the leading edge 68LE to the trailing edge68TE. A first width W1 of the restricted flow region 74-3 adjacent theleading edge 68LE is greater than a second width W2 of the restrictedflow region 74-3 adjacent the trailing edge 68TE such that at least amajority of the restricted flow region 74-3 progressively decreases inwidth or tapers from the leading edge 68LE to the trailing edge 68TE.The prioritized flow regions 74-1, 74-2 are axially aligned with therestricted flow region 74-3 for at least a majority, or more than 75% or90%, of a length of the liner segment 68 between the leading andtrailing edges 68LE, 68TE.

The perimeters P1, P2 of the prioritized flow region 74-1, 74-2 extendbetween the axially forwardmost and axially aftmost heat transferfeatures 72 that are along or otherwise near the respective nozzle axesN. A width of each of the prioritized flow regions 74-1, 74-2 cangenerally increase from the leading to trailing edges 68LE, 68TE. In theillustrated example of FIG. 6, a third width W3 of the first prioritizedflow region 74-1 adjacent the leading edge 68LE is less than a fourthwidth W4 of the first prioritized flow region 74-1 adjacent the trailingedge 68TE. A fifth width W5 of the second prioritized flow region 74-2adjacent the leading edge 68LE is less than a sixth width W6 of thesecond prioritized flow region 74-2 adjacent the trailing edge 68TE.Widths of the respective prioritized flow regions 74-1, 74-2 definingthe perimeters P1, P2 can progressively increase from the leading edge68LE to the trailing edge 68TE for at least a majority of theprioritized flow regions 74-1, 74-2.

The difference in widths of the flow regions 74 relative to the leadingand trailing edges 68LE, 68TE can increase diffusion of cooling flow CFejected from the trailing edge 68TE toward an adjacent, downstream linersegment 68-2 (shown in dashed lines for illustrated purposes). The heattransfer features 72 can be distributed such that the cooling flow CFejected along the trailing edge 68TE is diffused and substantiallyuniform in the circumferential direction when presented to the leadingedge 68LE of the downstream liner segment 68-2, which can reduce alikelihood of formation of hot spots along the downstream liner segment68-2.

During operation, cooling flow CF is communicated to each of the flowregions 74. The cooling flow CF can be communicated at substantially thesame temperature and/or pressure to each of the flow regions 74 adjacentto the leading edge 68LE, which serves as an inlet to the coolingcircuit 70. The cooling flow CF circulates across the heat transferfeatures 72 to provide convective cooling to adjacent portions of theexternal wall 68A.

The relative concentrations of the heat transfer features 72 in the flowregions 74 can cause at least a portion of the cooling flow CF in therestricted flow region 74-3 to be diverted or otherwise communicatedfrom the restricted flow region 74-3 to an adjacent one of theprioritized flow regions 74-1, 74-2 due to pressure gradient(s)established by the distribution of the heat transfer features 72-1,72-2, 72-3, with the prioritized flow regions 74-1, 74-2 operating atrelatively lower pressures. The distribution of heat transfer features72-3 establishes adverse pressure gradient(s) between the restrictedflow region 74-3 and prioritized flow regions 74-1, 74-2, which opposesmovement of the cooling flow CF from the prioritized flow regions 74-1,74-2 into the restricted flow region 74-3. At least some of the coolingflow CF can circulate through the restricted flow region 74-3 and isthen ejected from the restricted flow region 74-3 at the trailing edge68TE. The concentration of heat transfer features 72 in each of the flowregions 74 promotes communication of relatively more cooling flow CF inthe cooling circuit 70 along the nozzle axes N and toward the hotspot(s) HS, which can reduce a thermal gradient across the liner segment68 and improve durability of the combustor liner 60 (FIGS. 3 and 4).

FIG. 8 illustrates a combustor liner 160 according to another example.In this disclosure, like reference numerals designate like elementswhere appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding original elements. The combustor liner 160 includes aplurality of liner segments 168 including liner segment 168-1 arrangedcircumferentially adjacent to liner segments 168-2, 168-3 (shown indashed lines for illustrated purposes). Liner segment 168-1 defines acooling circuit 170 including first and second prioritized flow regions174-1, 174-2 that extend along opposed sides of restricted flow region174-3.

Each mate face 168M of the liner segment 168-1 is arranged to define anintersegment gap G with the mate faces 168M of adjacent liner segments168-2, 168-3. As illustrated by FIG. 3A, each intersegment gap G can bedimensioned to eject cooling flow CF radially inwardly or outwardly fromthe intersegment gap G into the combustor chamber 64 to provide coolingaugmentation to portions of the liner segments 68 adjacent the matefaces 68M.

Heat transfer features 172 are distributed in the cooling circuit 170 todefine second and third restricted flow regions 174-4, 174-5 includingrespective sets of the heat transfer features 172-4, 172-5. Each of thesecond and third restricted flow regions 174-4, 174-5 extends along arespective one of the mate faces 168M and bounds a perimeter of arespective one of the prioritized flow regions 174-1, 174-2. Each of theperimeters P1, P2 of the prioritized flow regions 174-1, 174-2 has asubstantially trapezoidal geometry, with a width of the prioritized flowregions 174-1, 174-2 generally increasing from leading edge 168LE totrailing edge 168TE. Perimeter P3 of the restricted flow region 174-3has a substantially trapezoidal geometry. Perimeters P4, P5 of therestricted flow regions 174-4, 174-5 each have a substantiallytriangular geometry. A width of each of the restricted flow regions174-4, 174-5 can be set (e.g., increased or decreased) to vary theamount of cooling flow CF communicated adjacent the mate faces 168M.

Each of the second and third restricted flow regions 174-4, 174-5 has arelatively greater concentration of the heat transfer features 172 thanan adjacent one of the prioritized flow regions 174-1, 174-2. Theprioritized flow regions 174-1, 174-2 have a relatively greater averageflow path volume than the restricted flow regions 174-3, 174-4, 174-5.The heat transfer features 172-4, 172-5 in the restricted flow regions174-4, 174-5 can oppose or otherwise reduce the amount of cooling flowCF that is communicated from the prioritized flow regions 174-1, 174-2toward the intersegment gaps G, which can reduce efficiency losses thatmay be otherwise caused by overcooling portions of the liner segments168 adjacent to the mate faces 168M. The distribution of heat transferfeatures 172-4, 172-5 can be the same or can differ from thedistribution of heat transfer features 172-3, including shape, spacingand/or orientation. In the illustrated example of FIG. 8, an averagesize of the heat transfer features 172-4, 172-5 is less than an averagesize of the heat transfer features 172-3, and an average spacing betweenthe heat transfer features 172-4, 172-5 is greater than an averagespacing between the heat transfer features 172-3.

FIG. 9 illustrates a liner segment 268 defining a cooling circuit 270according to yet another example. Heat transfer features 272-1, 272-4,272-5 are arranged to define restricted flow regions 274-4, 274-5 thatextend along mate faces 268M and along opposed sides of prioritized flowregion 274-1. Heat transfer features 272-1 have a relatively greaterdiameter and are less densely spaced than heat transfer features 274-4,274-5 such that the prioritized flow region 274-1 has a relativelygreater average flow path volume than the restricted flow regions 274-4,274-5 to promote flow of cooling flow CF toward and along a projectionof nozzle axis N. In the illustrated example of FIG. 10, heat transferfeatures 372-1 have a substantially rectangular geometry, with a majorcomponent of a length of the heat transfer features 372-1 oriented in acircumferential direction with respect to a respective nozzle axis Nand/or longitudinal axis X to oppose cooling flow CF.

FIG. 11 illustrates example geometries of heat transfer features 472that can be utilized in any of the cooling circuits disclosed herein.Heat transfer features 472A have a substantially cylindricalcross-sectional geometry. Heat transfer features 472B have anelliptical, non-circular geometry. Heat transfer features 472C have anelongated, substantially rectangular geometry. Heat transfer features472D have a diamond shaped geometry. Heat transfer features 472E have arace track shaped geometry. Heat transfer features 472F have a teardropshaped geometry. The arrangement of heat transfer features 472B, 472Ccan improve diffusion of cooling flow CF outwardly from trailing edge468TE of liner segment 468 and reduce distress by orienting the coolingflow CF toward a localized hot spot of a downstream liner segment.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A combustor liner for a gas turbine enginecomprising: at least one liner segment including an external walldimensioned to bound a combustion chamber, the external wall extendingbetween leading and trailing edges in an axial direction and extendingbetween opposed mate faces in a circumferential direction; and wherein acooling circuit is defined by the external wall, a plurality of heattransfer features are distributed in the cooling circuit to define afirst restricted flow region that tapers from the leading edge to thetrailing edge and to define at least one prioritized flow region thatextends substantially from the leading edge to the trailing edge suchthat the at least one prioritized flow region is bounded by a perimeterof the first restricted flow region, and the at least one prioritizedflow region has a lesser concentration of the plurality of heat transferfeatures than the first restricted flow region.
 2. The combustor lineras recited in claim 1, wherein the at least one prioritized flow regionincludes first and second prioritized flow regions on opposed sides ofthe first restricted flow region.
 3. The combustor liner as recited inclaim 2, wherein the plurality of heat transfer features are distributedin the cooling circuit to define second and third restricted flowregions that extend substantially along the mate faces to boundrespective ones of the first and second prioritized flow regions.
 4. Thecombustor liner as recited in claim 3, wherein each of the first andsecond prioritized flow regions has a substantially trapezoidalgeometry.
 5. The combustor liner as recited in claim 1, wherein the atleast one liner segment includes a first liner segment and a secondliner segment arranged in the axial direction to define a stepwisechange in area of the combustion chamber such that the cooling circuitof the first liner segment is oriented to eject cooling flow from thetrailing edge of the first liner segment onto external surfaces of theexternal wall of the second liner segment that defines the combustionchamber.
 6. The combustor liner as recited in claim 5, wherein the atleast one liner segment includes an array of liner segments, and each ofthe mate faces defines an intersegment gap with an adjacent one of theliner segments.
 7. The combustor liner as recited in claim 1, whereinthe plurality of heat transfer features includes a plurality ofpedestals that extend in a radial direction between opposed internalsurfaces defining the cooling circuit.
 8. The combustor liner as recitedin claim 1, wherein respective sets of the plurality of heat transferfeatures are uniformly distributed in the first restricted flow regionand in the at least one prioritized flow region.
 9. The combustor lineras recited in claim 1, further comprising a thermal barrier coatingdisposed on surfaces of the external wall defining the combustionchamber.
 10. The combustor liner as recited in claim 9, wherein thesurfaces of the external wall defining the combustion chamber aresubstantially free of any cooling apertures along the cooling circuit.11. The combustor liner as recited in claim 1, wherein the external wallis a bulkhead that bounds the combustion chamber in the axial direction,the bulkhead including at least one aperture along the combustionchamber that is dimensioned to receive a fuel injector nozzle.
 12. Acombustor section for a gas turbine engine comprising: an array of fuelinjector nozzles arranged about a longitudinal axis; a combustor linerincluding an array of liner segments arranged about the longitudinalaxis to define a combustion chamber; wherein each one of the fuelinjector nozzles defines a nozzle axis, a projection of the nozzle axisextending through the combustion chamber; wherein each one of the linersegments comprises: an external wall extending axially between leadingand trailing edges and extending circumferentially between opposed matefaces with respect to the longitudinal axis; wherein a cooling circuitis defined by the external wall; and wherein a plurality of heattransfer features are distributed in the cooling circuit to define firstand second prioritized flow regions on opposed sides of a firstrestricted flow region, each of the first and second prioritized flowregions extending axially along the projection of the nozzle axis ofrespective ones of the fuel injector nozzles from the leading edge tothe trailing edge such that the first restricted flow region tapers fromthe leading edge to the trailing edge, and each of the first and secondprioritized flow regions having a relatively greater average flow pathvolume than the first restricted flow region.
 13. The combustor sectionas recited in claim 12, wherein the plurality of heat transfer featuresare distributed in the cooling circuit to define second and thirdrestricted flow regions that extend substantially along the mate facesto bound a perimeter of respective ones of the first and secondprioritized flow regions, and each of the first and second prioritizedflow regions has a relatively greater average flow path volume than thesecond and third restricted flow regions.
 14. The combustor section asrecited in claim 12, wherein surfaces of the external wall defining thecombustion chamber are substantially free of any cooling apertures alongthe cooling circuit.
 15. A gas turbine engine comprising: a compressorsection; a turbine section that drives the compressor section; acombustor section including a combustor, the combustor including acombustor liner and an array of fuel injector nozzles arranged about anengine longitudinal axis; wherein the combustor liner includes an arrayof liner segments arranged about the engine longitudinal axis to definea combustion chamber; wherein each one of the fuel injector nozzlesdefines a nozzle axis, a projection of the nozzle axis extending throughthe combustion chamber; and wherein each one of the liner segmentscomprises: an external wall extending axially between leading andtrailing edges and extending circumferentially between opposed matefaces with respect to the engine longitudinal axis; wherein a coolingcircuit is defined by the external wall; and wherein a plurality of heattransfer features are distributed in the cooling circuit to define firstand second prioritized flow regions on opposed sides of a firstrestricted flow region, each of the first and second prioritized flowregions extending axially along the projection of the nozzle axis ofrespective ones of the fuel injector nozzles such that a width of thefirst and second prioritized flow regions progressively increases fromthe leading edge to the trailing edge.
 16. The gas turbine engine asrecited in claim 15, wherein the first restricted flow region iscircumferentially spaced from the projection of the nozzle axis of eachand every one of the fuel injector nozzles.
 17. The gas turbine engineas recited in claim 15, wherein the first restricted flow region tapersfrom the leading edge to the trailing edge.
 18. The gas turbine engineas recited in claim 15, wherein: each of the mate faces defines anintersegment gap with the mate face of an adjacent one of the linersegments; and the plurality of heat transfer features are distributed inthe cooling circuit to define second and third restricted flow regionsthat extend substantially along the mate faces to bound a perimeter ofrespective ones of the first and second prioritized flow regions. 19.The gas turbine engine as recited in claim 18, wherein the intersegmentgap is dimensioned to eject cooling flow into the combustion chamber.20. The gas turbine engine as recited in claim 15, wherein the array ofliner segments includes a first set of liner segments and a second setof liner segments axially arranged to define a stepwise change in areaof the combustion chamber such that each cooling circuit of the firstset of liner segments is oriented to eject cooling flow onto externalsurfaces of the second set of liner segments bounding the combustionchamber.